Multiple staged compressor with last stage airfoil cooling

ABSTRACT

A high compression ratio compressor having multiple stages of airfoils to produce the high pressure rations, where the last stage airfoils are cooled by drawing heat away from the airfoils using a heat pipe, and then passing cooling air through a heat exchanger associated with the heat pipe to draw heat away from the heat pipe. The cooling air used for cooling the airfoils is bled off from an earlier stage than the last stage and then discharged into a stage upstream from the bleed off location.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a CONTINUATION-IN-PART of U.S. patent applicationSer. No. 12/268,340, now U.S. Pat. No. 8,240,975, filed on Nov. 10, 2008and entitled MULTIPLE STAGED COMPRESSOR WITH LAST STAGE AIRFOIL COOLING.

GOVERNMENT LICENSE RIGHTS

None.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates go gas turbine engines, and morespecifically to a high pressure ratio compressor with last stage airfoilcooling used in a gas turbine engine.

2. Description of the Related Art Including Information Disclosed Under37 CFR 1.97 and 1.98

A gas turbine engine, such as an aero engine used to power an aircraftor an industrial gas turbine engine used to produce electric power, is avery efficient power plant. The compressed air from a compressor ispassed into a combustor where the air is burned with a fuel to produce ahot gas flow, the hot gas flow is then passed through a turbine to drivethe rotor shaft connected to the compressor and, in the case of an aeroengine produce thrust and/or drive the fan, or in the case of the IGTdrive an electric generator to produce the electric power. In bothcases, the efficiency of the engine can be increased by passing a highertemperature gas into the turbine.

Modern gas turbine engines have multiple stages in the compressor inorder to produce the very high pressure ratios between the outlet of thecompressor and the inlet. For example, the Pratt & Whitney F100 aeroengine that is used to power the military F15 and F16 fighter aircraftincludes 13 stages in the compressor and produces a pressure ratio of 30to 1 (the outlet pressure is 30 times the inlet pressure). A higherpressure ratio will allow for higher efficiencies for the engine. Withthe recent improvements in compressor design, a higher number of stagescan be used to produce an even higher pressure ratio. Futures aeroengines are anticipated to have compressor ratios in the 50s.

However, as the air through the compressor is compressed, thetemperature of the compressed air also increases. A multiple stagecompressor will generally add 90 degrees F. to the compressed air foreach stage. As the number of stages in the compressor grows, thecompressor outlet air becomes higher to the point where the last stageairfoils (stator vanes and compressor blades) become so hot that theairfoils can be damaged from the high thermal load. Thus, there is aneed in the prior art for a multiple stage compressor with a very highpressure ratio to have cooling of the last stage airfoils in thecompressor in order to withstand the higher temperatures.

BRIEF SUMMARY OF THE INVENTION

A compressor in a gas turbine engine in which the compression ratio isso high that the last stage rotor blades and stator vanes requirecooling to prevent overheating of these airfoils due to the higher airtemperature resulting from the increased compression ratio produced bythe engine. The last stage blades and vanes include internal cooling airpassages to produce cooling through convection and impingement. Thecompressed air used for cooling these airfoils is bled off from anupstream stage of the compressor, passed through the airfoils forcooling, and then reintroduced into the compressor at an upper stagefrom where the bleed off air was first extracted.

Because the air pressure used for the cooling of the last stage airfoilsis less than the external air flow pressure around these cooled airfoils, the cooled air foils cannot include discharging film cooling airfor cooling the exterior surface because of the differential pressure.The cooled airfoils are thus cooled by a closed system and the spentcooling air reintroduced into the compressor at a location upstream fromthe bleed off location.

In a second embodiment of the present invention, heat pipes located inthe last stage or later stages airfoils are used to draw heat away fromthe airfoils, and the cooling air passes through heat exchangesassociated with the heat pipes to remove heat and cool the airfoils.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

FIG. 1 shows a cross section view of a first embodiment of the multiplestage compressor of the present invention.

FIG. 2 shows a cross section view of a second embodiment of the multiplestage compressor of the present invention having heat pipes.

DETAILED DESCRIPTION OF THE INVENTION

The present invention is a multiple stage compressor that produces avery high pressure ratio such that cooling of the last stage or stagesof the compressor air required. The compressor is intended to be used ina gas turbine engine such as an aero engine or an IGT engine. However,the present invention could be used in any turbomachine in which amultiple stage compressor is used that produces the high pressure ratioin which cooling of the last stage airfoils is required to preventthermal damage.

FIG. 1 shows a cross section of the compressor of the present inventionin which a number of stages are present with each stage having a statoror guide vane located upstream from an associated rotor blade. In atypical multiple stage compressor of an aero engine, the outer diameterof the compressor is about at a constant radial diameter while the innerdiameter is conical shaped with a decreasing airfoil spanwise height inthe downstream direction. The inlet air to the compressor is atatmospheric pressure. The compressor progressively compresses the air asthe compressed air passes through the multiple stages. As the air iscompressed, the temperature of the compressed air increases. A typicalcompressor will increase the compressed air temperature about 90 degreesF. in each stage. In the last stages, the compressed air can be at sucha high temperature that the airfoils can be damaged from the hightemperature. The material properties of these airfoils are such that thehigh temperature compressed airs passing through these airfoils exceedthe safe temperature level for the materials.

In the present invention is FIG. 1, the last stage vanes 12 and blades11 include closed loop cooling air passages to pass cooling air throughthe airfoils without discharging the cooling air into the hightemperature compressed air stream through the compressor. The internalairfoil cooling passages can be any type of prior art closed loopcooling passage circuit that makes use of well known convection coolingand impingement cooling of airfoils. Compressed cooling air from amiddle stage 15 of the compressor is bled off and passed through acooling air passage in the rotor shaft assembly and into the inlet ofthe internal cooling air passage of the rotor blade. The cooling airpasses through the rotor blade cooling passage and then flows through areturn air passage also in the rotor shaft to be discharged into thecompressor at a stage 16 upstream from the bled off stage. This is dueto the loss of pressure in the cooling air from passing through thecooling passages in the rotor shaft and the rotor blade.

To cool the last stage stator vane 11, cooling air is also bled off fromthe compressor at a middle stage 13 and directed through a cooling airpassage and into the internal cooling air passage formed within thestator vane. The cooling air passes through the vane cooling passage,and is then directed through a return air cooling air passage and intothe compressor at a stage 14 upstream from the bled off location. Thisis also due to the loss of pressure in the cooling air from passingthrough the cooling supply passages in the casing and the stator vane.The bleed off air used for cooling of the last stage airfoils is fromthe lowest stage that would produce enough pressure to pass through thecooling circuit for the airfoils while still allowing for the spentcooling air to be discharged into an upstream stage of the compressor.The further down the compressor stages that the cooling air is bled offfrom, the higher the temperature of the cooling air used to pass throughthe airfoils for cooling.

Bleeding off the compressed air used for the cooling and thenre-supplying the cooling air back into the compressor minimizes the lossin the compressor. The heat picked up from the cooling air passingthrough the cooling passages within the airfoils is passed back into thecompressor mainstream air. The only significant losses are due to thepressure loss from the cooling air passing through the cooling passagesfrom the bleed off location to the re-supply location.

In other embodiments, other stages of the blades and vanes in thecompressor can also be cooled by passing bleed off cooling air throughthe internal cooling passages and then re-supplying the cooling air tothe compressor. The number of stages in the compressor that requirecooling would depend upon the compressed air temperature passing throughthose stages. Also, the stage at which the cooling air is bled off willdepend upon the required pressure for the cooling air that is needed topass through the cooling air passages and be discharged back into thecompressor. The re-supply locations will depend upon the pressuredifference between the main stream compressed air and the re-supplycooling air. The re-supply cooling air must be at a higher pressure thanthe mainstream compressor air or a backflow will occur. Since theairfoil internal cooling passage is a closed loop passage (no coolingair is discharged from the airfoil out into the mainstream compressorair flow), the pressure of the cooling air can be lower than thepressure of the main stream compressed air passing through that airfoil.

Also, in another embodiment, the cooling air can be discharged into theturbine section to provide cooling for turbine airfoils such as rotorblades and stator vanes and then discharged into the hot gas flowpassing through the turbine if the pressure differential is high enoughto prevent backflow into the turbine airfoils.

In another embodiment, the cooling air from the compressor airfoils canbe passed through a turbocharger to increase the pressure of the coolingair, and then passed into the combustor to be burned with the fuel. Withthis embodiment, the heated cooling air is burned with the fuel toproduce the hot gas flow that is passed through the turbine to drive therotor shaft.

FIG. 2 shows an embodiment of the present invention in which the coolingair does not pass through the last stage airfoils. In the closed loopcooling circuit of FIG. 1, if an airfoil was to crack then the hotcompressed air from the compressor can leak into the internal coolingpassages of the last stage airfoils. To prevent this, the FIG. 2embodiment uses heat pipes 21 and 23 that extend into the last stageairfoils 11 and 12 to draw heat away and into heat exchangers 22 and 24.The compressed air bled off from the compressor at 13 and 15 is passedthrough the heat exchangers 22 and 24 to draw heat away from the heatpipes 21 and 23 and thus the last stage airfoils 11 and 12 to cool theairfoils. With the use of heat pipes in the last stage airfoils 11 and12, no high temperature compressed air can leak into the cooling airpassages.

We claim:
 1. A multiple stage compressor comprising: a plurality ofstages of airfoils; the last stage of airfoils having heat pipe toprovide cooling for the airfoil; a heat pipe extending into the laststage airfoil to draw heat away from the last stage airfoil; a heatexchanger associated with the heat pipe to draw heat away from the heatpipe; a cooling air supply passage to supply cooling air to the heatexchanger; the cooling air supply passage connected to a stage of thecompressor having a lower pressure than the last stage airfoil; acooling air discharge passage to discharge spent cooling air from theheat exchanger of the last stage airfoil; and, the cooling air dischargepassage connected to a different stage of the compressor having a lowerpressure than the stage supplying the compressor cooling air to the laststage airfoil.
 2. The multiple stage compressor of claim 1, and furthercomprising: the last stage airfoils are rotor blades or stator vanes. 3.The multiple stage compressor of claim 1, and further comprising: thecompressor is an axial flow compressor used in a gas turbine engine. 4.The multiple stage compressor of claim 1, and further comprising: thecooling circuit in the last stage airfoils is a closed cooling circuitsuch that cooling air is not discharged out from the last stage airfoilsas film cooling air.
 5. The multiple stage compressor of claim 2, andfurther comprising: the stage that supplies cooling air to the airfoilsis the same; and, the stage that discharges cooling air from theairfoils is the same.
 6. The multiple stage compressor of claim 1, andfurther comprising: the compressor is a high compression ratiocompressor.
 7. A process for cooling a last stage airfoil of a highcompression ratio compressor, the compressor having a plurality ofstages, the process comprising the steps of: bleeding off a portion ofthe compressed air from the compressor at a stage that producescompressed air at a first pressure; drawing heat away from the laststage airfoil and into a heat exchanger; passing the compressed air fromthe bleed off air through the heat exchanger to provide cooling of thelast stage airfoil; and, discharging the heated cooling air from theheat exchanger into the compressor at a stage that produces compressedair at a second pressure that is lower than the first pressure.
 8. Theprocess for cooling a last stage airfoil of a high compression ratiocompressor of claim 7, and further comprising the step of: drawing heataway from the last stage rotor blades.
 9. The process for cooling a laststage airfoil of a high compression ratio compressor of claim 7, andfurther comprising the step of: drawing heat away from the last stagestator vanes and the last stage rotor blades.
 10. The process forcooling a last stage airfoil of a high compression ratio compressor ofclaim 7, and further comprising the step of: locating the heat exchangerfor the last stage rotor blades within a rotor shaft of the compressor;and, passing the cooling air through a passage located within the rotorshaft to draw heat away from the heat exchanger.